Early turbojet engines were not very fuel-efficient as their overall pressure ratio and turbine inlet temperature were severely limited by the technology available at the time. In 1939-1941 Soviet designer Arkhip Lyulka elaborated the design for the world's first turbofan engine, and acquired a patent for this new invention on April 22, 1941. Although several prototypes were built and ready for testing, Lyulka was in 1941 forced to abandon his research and evacuate to the Urals following the Nazi invasion of the Soviet Union. So the first turbofan to run was apparently the German Daimler-Benz DB 670 (designated as the 109-007 by the RLM) with a first run date of 27 May 1943. Turbomachinery testing, using an electric motor, had started on 1 April 1943. The engine was abandoned later while the war went on and problems could not be solved. The British wartime Metrovick F.2 axial flow jet was given a fan, as the Metrovick F.3 in 1943, to create the first British turbofan.
Improved materials, and the introduction of twin compressors such as in the Bristol Olympus and Pratt & Whitney JT3C engines, increased the overall pressure ratio and thus the thermodynamic efficiency of engines, but they also led to a poor propulsive efficiency, as pure turbojets have a high specific thrust/high velocity exhaust better suited to supersonic flight.
The original low-bypass turbofan engines were designed to improve propulsive efficiency by reducing the exhaust velocity to a value closer to that of the aircraft. The Rolls-Royce Conway, the world's first production turbofan, had a bypass ratio of 0.3, similar to the modern General Electric F404 fighter engine. Civilian turbofan engines of the 1960s, such as the Pratt & Whitney JT8D and the Rolls-Royce Spey had bypass ratios closer to 1, and were similar to their military equivalents.
The first General Electric turbofan was the aft-fan CJ805-23 based on the CJ805-3 turbojet. It was followed by the aft-fan General Electric CF700 engine with a 2.0 bypass ratio. This was derived from the General Electric J85/CJ610 turbojet (2,850 lbf or 12,650 N) to power the larger Rockwell Sabreliner 75/80 model aircraft, as well as the Dassault Falcon 20 with about a 50% increase in thrust (4,200 lbf or 18,700 N). The CF700 was the first small turbofan in the world to be certified by the Federal Aviation Administration (FAA). There were at one time over 400 CF700 aircraft in operation around the world, with an experience base of over 10 million service hours. The CF700 turbofan engine was also used to train Moon-bound astronauts in Project Apollo as the powerplant for the Lunar Landing Research Vehicle.
A high-specific-thrust/low-bypass-ratio turbofan normally has a multi-stage fan, developing a relatively high pressure ratio and, thus, yielding a high (mixed or cold) exhaust velocity. The core airflow needs to be large enough to give sufficient core power to drive the fan. A smaller core flow/higher bypass ratio cycle can be achieved by raising the (HP) turbine rotor inlet temperature.
To illustrate one aspect of how a turbofan differs from a turbojet they may be compared, as in a re-engining assessment, at the same airflow (to keep a common intake for example) and the same net thrust (i.e. same specific thrust). A bypass flow can only be added to the turbojet if the turbine inlet temperature is allowed to increase to compensate for the smaller core flow. Improvements in turbine cooling/material technology would allow the use of a higher turbine inlet temperature despite an increase in cooling air temperature which would result from any overall pressure-ratio increase.
The resulting turbofan, with reasonable efficiencies and duct loss for the added components, would probably operate at a higher nozzle pressure ratio than the turbojet, but with a lower exhaust temperature to retain net thrust. Since the temperature rise across the whole engine (intake to nozzle) would be lower, the (dry power) fuel flow would also be reduced, resulting in a better specific fuel consumption (SFC).
Some low-bypass ratio military turbofans (e.g. F404) have variable inlet guide vanes to direct air onto the first fan rotor stage. This improves the fan surge margin (see compressor map).
Since the 1970s, most jet fighter engines have been low/medium bypass turbofans with a mixed exhaust, afterburner and variable area final nozzle. An afterburner is a combustor located downstream of the turbine blades and directly upstream of the nozzle, which burns fuel from afterburner-specific fuel injectors. When lit, prodigious amounts of fuel are burnt in the afterburner, raising the temperature of exhaust gases by a significant degree, resulting in a higher exhaust velocity/engine specific thrust. The variable geometry nozzle must open to a larger throat area to accommodate the extra volume flow when the afterburner is lit. Afterburning is often designed to give a significant thrust boost for take off, transonic acceleration and combat maneuvers, but is very fuel intensive. Consequently, afterburning can only be used for short portions of a mission.
Unlike the main combustor, where the downstream turbine blades must not be damaged by high temperatures, an afterburner can operate at the ideal maximum (stoichiometric) temperature (i.e., about 2100K/3780Ra/3320F). At a fixed total applied fuel:air ratio, the total fuel flow for a given fan airflow will be the same, regardless of the dry specific thrust of the engine. However, a high specific thrust turbofan will, by definition, have a higher nozzle pressure ratio, resulting in a higher afterburning net thrust and, therefore, a lower afterburning specific fuel consumption (SFC). However, high specific thrust engines have a high dry SFC. The situation is reversed for a medium specific thrust afterburning turbofan: i.e., poor afterburning SFC/good dry SFC. The former engine is suitable for a combat aircraft which must remain in afterburning combat for a fairly long period, but only has to fight fairly close to the airfield (e.g. cross border skirmishes) The latter engine is better for an aircraft that has to fly some distance, or loiter for a long time, before going into combat. However, the pilot can only afford to stay in afterburning for a short period, before aircraft fuel reserves become dangerously low.
The first production afterburning turbofan engine was the Pratt & Whitney TF30, which initially powered the F-111 Aardvark and F-14 Tomcat. Current low-bypass military turbofans include the Pratt & Whitney F119, the Eurojet EJ200, the General Electric F110, the Klimov RD-33, and the Saturn AL-31, all of which feature a mixed exhaust, afterburner and variable area propelling nozzle.
The low-specific-thrust/high-bypass-ratio turbofans used in today's civil jetliners (and some military transport aircraft) evolved from the high-specific-thrust/low-bypass-ratio turbofans used in such aircraft back in the 1960s.
Low specific thrust is achieved by replacing the multi-stage fan with a single-stage unit. Unlike some military engines, modern civil turbofans do not have any stationary inlet guide vanes in front of the fan rotor. The fan is scaled to achieve the desired net thrust.
The core (or gas generator) of the engine must generate sufficient core power to at least drive the fan at its design flow and pressure ratio. Through improvements in turbine cooling/material technology, a higher (HP) turbine rotor inlet temperature can be used, thus facilitating a smaller (and lighter) core and (potentially) improving the core thermal efficiency. Reducing the core mass flow tends to increase the load on the LP turbine, so this unit may require additional stages to reduce the average stage loading and to maintain LP turbine efficiency. Reducing core flow also increases bypass ratio. Bypass ratios greater than 5:1 are increasingly common with the Pratt & Whitney PW1000G attaining 12.5:1.
Further improvements in core thermal efficiency can be achieved by raising the overall pressure ratio of the core. Improved blade aerodynamics reduces the number of extra compressor stages required. With multiple compressors (i.e., LPC, IPC, and HPC) dramatic increases in overall pressure ratio have become possible. Variable geometry (i.e., stators) enable high-pressure-ratio compressors to work surge-free at all throttle settings.
The first (experimental) high-bypass turbofan engine was built and run on February 13, 1964 by AVCO-Lycoming. Shortly after, the General Electric TF39 became the first production model, designed to power the Lockheed C-5 Galaxy military transport aircraft. The civil General Electric CF6 engine used a derived design. Other high-bypass turbofans are the Pratt & Whitney JT9D, the three-shaft Rolls-Royce RB211 and the CFM International CFM56; also the smaller TF34. More recent large high-bypass turbofans include the Pratt & Whitney PW4000, the three-shaft Rolls-Royce Trent, the General Electric GE90/GEnx and the GP7000, produced jointly by GE and P&W.
For reasons of fuel economy, and also of reduced noise, almost all of today's jet airliners are powered by high-bypass turbofans. Although modern combat aircraft tend to use low-bypass ratio turbofans, military transport aircraft (e.g., C-17 ) mainly use high-bypass ratio turbofans (or turboprops) for fuel efficiency.
The lower the specific thrust of a turbofan, the lower the mean jet outlet velocity, which in turn translates into a high thrust lapse rate ( i.e. decreasing thrust with increasing flight speed). See technical discussion below, item 2. Consequently, an engine sized to propel an aircraft at high subsonic flight speed (e.g., Mach 0.83) generate a relatively high thrust at low flight speed, thus enhancing runway performance. Low specific thrust engines tend to have a high bypass ratio, but this is also a function of the temperature of the turbine system.
The turbofans on twin engined airliners are further more powerful to cope with losing one engine during take-off, which reduces the aircraft's net thrust by half. Modern twin engined airliners normally climb very steeply immediately after take-off. If one engine is lost, the climb-out is much shallower, but sufficient to clear obstacles in the flightpath.
The Soviet Union's engine technology was less advanced than the West's and its first wide-body aircraft, the Ilyushin Il-86, was powered by low-bypass engines. The Yakovlev Yak-42, a medium-range, rear-engined aircraft seating up to 120 passengers introduced in 1980 was the first Soviet aircraft to use high-bypass engines.
Turbofan engines come in a variety of engine configurations. For a given engine cycle (i.e., same airflow, bypass ratio, fan pressure ratio, overall pressure ratio and HP turbine rotor inlet temperature), the choice of turbofan configuration has little impact upon the design point performance (e.g., net thrust, SFC), as long as overall component performance is maintained. Off-design performance and stability is, however, affected by engine configuration.
As the design overall pressure ratio of an engine cycle increases, it becomes more difficult to operate at low rpm, without encountering an instability known as compressor surge. This occurs when some of the compressor aerofoils stall (like the wings of an aircraft) causing a violent change in the direction of the airflow. However, compressor stall can be avoided, at low rpm, by progressively:
- opening interstage/intercompressor blow-off valves (inefficient), and/or
- closing variable stators within the compressor
Most modern American civil turbofans employ a relatively high-pressure-ratio high-pressure (HP) compressor, with many rows of variable stators to control surge margin at low rpm. In the three-spool RB211/Trent the core compression system is split into two, with the IP compressor, which supercharges the HP compressor, being on a different coaxial shaft and driven by a separate (IP) turbine. As the HP compressor has a modest pressure ratio its speed can be reduced surge-free, without employing variable geometry. However, because a shallow IP compressor working line is inevitable, the IPC has one stage of variable geometry on all variants except the -535, which has none.
Although far from common, the single-shaft turbofan is probably the simplest configuration, comprising a fan and high-pressure compressor driven by a single turbine unit, all on the same shaft. The SNECMA M53, which powers Mirage fighter aircraft, is an example of a single-shaft turbofan. Despite the simplicity of the turbomachinery configuration, the M53 requires a variable area mixer to facilitate part-throttle operation.
One of the earliest turbofans was a derivative of the General Electric J79 turbojet, known as the CJ805-23, which featured an integrated aft fan/low-pressure (LP) turbine unit located in the turbojet exhaust jetpipe. Hot gas from the turbojet turbine exhaust expanded through the LP turbine, the fan blades being a radial extension of the turbine blades. This aft-fan configuration was later exploited in the General Electric GE-36 UDF (propfan) demonstrator of the early 80s. One of the problems with the aft fan configuration is hot gas leakage from the LP turbine to the fan.
Many turbofans have the basic two-spool configuration where both the fan and LP turbine (i.e., LP spool) are mounted on a second (LP) shaft, running concentrically with the HP spool (i.e., HP compressor driven by HP turbine). The BR710 is typical of this configuration. At the smaller thrust sizes, instead of all-axial blading, the HP compressor configuration may be axial-centrifugal (e.g., General Electric CFE738), double-centrifugal or even diagonal/centrifugal (e.g., Pratt & Whitney Canada PW600).
Higher overall pressure ratios can be achieved by either raising the HP compressor pressure ratio or adding an intermediate-pressure (IP) compressor between the fan and HP compressor, to supercharge or boost the latter unit helping to raise the overall pressure ratio of the engine cycle to the very high levels employed today (i.e., greater than 40:1, typically). All of the large American turbofans (e.g., General Electric CF6, GE90 and GEnx plus Pratt & Whitney JT9D and PW4000) feature an IP compressor mounted on the LP shaft and driven, like the fan, by the LP turbine, the mechanical speed of which is dictated by the tip speed and diameter of the fan. The Rolls-Royce BR715 is a non-American example of this. The high bypass ratios (i.e., fan duct flow/core flow) used in modern civil turbofans tends to reduce the relative diameter of the attached IP compressor, causing its mean tip speed to decrease. Consequently, more IPC stages are required to develop the necessary IPC pressure rise.
Rolls-Royce chose a three-spool configuration for their large civil turbofans (i.e., the RB211 and Trent families), where the intermediate pressure (IP) compressor is mounted on a separate (IP) shaft, running concentrically with the LP and HP shafts, and is driven by a separate IP turbine. The first three-spool engine was the earlier Rolls-Royce RB.203 Trent of 1967.
Ivchenko Design Bureau chose the same configuration for their Lotarev D-36 engine, followed by Lotarev/Progress D-18T and Progress D-436.
The Turbo-Union RB199 military turbofan also has a three-spool configuration, as do the military Kuznetsov NK-25 and NK-321.
As bypass ratio increases, the mean radius ratio of the fan and low-pressure turbine (LPT) increases. Consequently, if the fan is to rotate at its optimum blade speed the LPT blading will spin slowly, so additional LPT stages will be required, to extract sufficient energy to drive the fan. Introducing a (planetary) reduction gearbox, with a suitable gear ratio, between the LP shaft and the fan enables both the fan and LP turbine to operate at their optimum speeds. Typical of this configuration are the long-established Honeywell TFE731, the Honeywell ALF 502/507, and the recent Pratt & Whitney PW1000G.
Most of the configurations discussed above are used in civilian turbofans, while modern military turbofans (e.g., SNECMA M88) are usually basic two-spool.
Most civil turbofans use a high-efficiency, 2-stage HP turbine to drive the HP compressor. The CFM56 uses an alternative approach: a single-stage, high-work unit. While this approach is probably less efficient, there are savings on cooling air, weight and cost.
In the RB211 and Trent 3-spool engine series, the HP compressor pressure ratio is modest so only a single HP stage is required. Rather than adding stage/s to the LP turbine to drive the higher pressure ratio IP (intermediate pressure) compressor, Rolls-Royce mounts it on a separate shaft and drives it with an IP turbine.
Because the HP compressor pressure ratio is modest, modern military turbofans tend to use a single-stage HP turbine.
Modern civil turbofans have multi-stage LP turbines (e.g., 3, 4, 5, 6, 7). The number of stages required depends on the engine cycle bypass ratio and how much supercharging (i.e., IP compression) is on the LP shaft, behind the fan. A geared fan may reduce the number of required LPT stages in some applications. Because of the much lower bypass ratios employed, military turbofans only require one or two LP turbine stages.
Consider a mixed turbofan with a fixed bypass ratio and airflow. Increasing the overall pressure ratio of the compression system raises the combustor entry temperature. Therefore, at a fixed fuel flow there is an increase in (HP) turbine rotor inlet temperature. Although the higher temperature rise across the compression system implies a larger temperature drop over the turbine system, the mixed nozzle temperature is unaffected, because the same amount of heat is being added to the system. There is, however, a rise in nozzle pressure, because overall pressure ratio increases faster than the turbine expansion ratio, causing an increase in the hot mixer entry pressure. Consequently, net thrust increases, whilst specific fuel consumption (fuel flow/net thrust) decreases. A similar trend occurs with unmixed turbofans.
So turbofans can be made more fuel efficient by raising overall pressure ratio and turbine rotor inlet temperature in unison. However, better turbine materials and/or improved vane/blade cooling are required to cope with increases in both turbine rotor inlet temperature and compressor delivery temperature. Increasing the latter may require better compressor materials.
Overall pressure ratio can be increased by improving fan (or) LP compressor pressure ratio and/or HP compressor pressure ratio. If the latter is held constant, the increase in (HP) compressor delivery temperature (from raising overall pressure ratio) implies an increase in HP mechanical speed. However, stressing considerations might limit this parameter, implying, despite an increase in overall pressure ratio, a reduction in HP compressor pressure ratio.
According to simple theory, if the ratio of turbine rotor inlet temperature/(HP) compressor delivery temperature is maintained, the HP turbine throat area can be retained. However, this assumes that cycle improvements are obtained, while retaining the datum (HP) compressor exit flow function (non-dimensional flow). In practice, changes to the non-dimensional speed of the (HP) compressor and cooling bleed extraction would probably make this assumption invalid, making some adjustment to HP turbine throat area unavoidable. This means the HP turbine nozzle guide vanes would have to be different from the original. In all probability, the downstream LP turbine nozzle guide vanes would have to be changed anyway.
Thrust growth is obtained by increasing core power. There are two basic routes available:
- hot route: increase HP turbine rotor inlet temperature
- cold route: increase core mass flow
Both routes require an increase in the combustor fuel flow and, therefore, the heat energy added to the core stream.
The hot route may require changes in turbine blade/vane materials and/or better blade/vane cooling. The cold route can be obtained by one of the following:
- adding T-stages to the LP/IP compression
- adding a zero-stage to the HP compression
- improving the compression process, without adding stages (e.g. higher fan hub pressure ratio)
all of which increase both overall pressure ratio and core airflow.
Alternatively, the core size can be increased, to raise core airflow, without changing overall pressure ratio. This route is expensive, since a new (upflowed) turbine system (and possibly a larger IP compressor) is also required.
Changes must also be made to the fan to absorb the extra core power. On a civil engine, jet noise considerations mean that any significant increase in take-off thrust must be accompanied by a corresponding increase in fan mass flow (to maintain a T/O specific thrust of about 30 lbf/lb/s).
- Specific thrust (net thrust/intake airflow) is an important parameter for turbofans and jet engines in general. Imagine a fan (driven by an appropriately sized electric motor) operating within a pipe, which is connected to a propelling nozzle. It is fairly obvious, the higher the fan pressure ratio (fan discharge pressure/fan inlet pressure), the higher the jet velocity and the corresponding specific thrust. Now imagine we replace this set-up with an equivalent turbofan - same airflow and same fan pressure ratio. Obviously, the core of the turbofan must produce sufficient power to drive the fan via the low-pressure (LP) turbine. If we choose a low (HP) turbine inlet temperature for the gas generator, the core airflow needs to be relatively high to compensate. The corresponding bypass ratio is therefore relatively low. If we raise the turbine inlet temperature, the core airflow can be smaller, thus increasing bypass ratio. Raising turbine inlet temperature tends to increase thermal efficiency and, therefore, improve fuel efficiency.
- Naturally, as altitude increases, there is a decrease in air density and, therefore, the net thrust of an engine. There is also a flight speed effect, termed thrust lapse rate. Consider the approximate equation for net thrust again:
With a high specific thrust (e.g., fighter) engine, the jet velocity is relatively high, so intuitively one can see that increases in flight velocity have less of an impact upon net thrust than a medium specific thrust (e.g., trainer) engine, where the jet velocity is lower. The impact of thrust lapse rate upon a low specific thrust (e.g., civil) engine is even more severe. At high flight speeds, high-specific-thrust engines can pick up net thrust through the ram rise in the intake, but this effect tends to diminish at supersonic speeds because of shock wave losses.
- Thrust growth on civil turbofans is usually obtained by increasing fan airflow, thus preventing the jet noise becoming too high. However, the larger fan airflow requires more power from the core. This can be achieved by raising the overall pressure ratio (combustor inlet pressure/intake delivery pressure) to induce more airflow into the core and by increasing turbine inlet temperature. Together, these parameters tend to increase core thermal efficiency and improve fuel efficiency.
- Some high-bypass-ratio civil turbofans use an extremely low area ratio (less than 1.01), convergent-divergent, nozzle on the bypass (or mixed exhaust) stream, to control the fan working line. The nozzle acts as if it has variable geometry. At low flight speeds the nozzle is unchoked (less than a Mach number of unity), so the exhaust gas speeds up as it approaches the throat and then slows down slightly as it reaches the divergent section. Consequently, the nozzle exit area controls the fan match and, being larger than the throat, pulls the fan working line slightly away from surge. At higher flight speeds, the ram rise in the intake increases nozzle pressure ratio to the point where the throat becomes choked (M=1.0). Under these circumstances, the throat area dictates the fan match and, being smaller than the exit, pushes the fan working line slightly towards surge. This is not a problem, since fan surge margin is much better at high flight speeds.
- The off-design behaviour of turbofans is illustrated under compressor map and turbine map.
- Because modern civil turbofans operate at low specific thrust, they only require a single fan stage to develop the required fan pressure ratio. The desired overall pressure ratio for the engine cycle is usually achieved by multiple axial stages on the core compression. Rolls-Royce tend to split the core compression into two with an intermediate pressure (IP) supercharging the HP compressor, both units being driven by turbines with a single stage, mounted on separate shafts. Consequently, the HP compressor need only develop a modest pressure ratio (e.g., ~4.5:1). US civil engines use much higher HP compressor pressure ratios (e.g., ~23:1 on the General Electric GE90) and tend to be driven by a two-stage HP turbine. Even so, there are usually a few IP axial stages mounted on the LP shaft, behind the fan, to further supercharge the core compression system. Civil engines have multi-stage LP turbines, the number of stages being determined by the bypass ratio, the amount of IP compression on the LP shaft and the LP turbine blade speed.
- Because military engines usually have to be able to fly very fast at sea level, the limit on HP compressor delivery temperature is reached at a fairly modest design overall pressure ratio, compared with that of a civil engine. Also the fan pressure ratio is relatively high, to achieve a medium to high specific thrust. Consequently, modern military turbofans usually only have 5 or 6 HP compressor stages and only require a single-stage HP turbine. Low-bypass-ratio military turbofans usually have one LP turbine stage, but higher bypass ratio engines need two stages. In theory, by adding IP compressor stages, a modern military turbofan HP compressor could be used in a civil turbofan derivative, but the core would tend to be too small for high thrust applications.
Modern commercial aircraft employ high-bypass-ratio (HBPR) engines with separate flow, non-mixing, short-duct exhaust systems. These propulsion systems are known to generate significantly high noise levels due to the high-speed, high-temperature, and high-pressure nature of the exhaust jet, especially during high thrust conditions such as those required for takeoff. The primary source of jet noise is the turbulent mixing of shear layers in the engine’s exhaust. These shear layers contain instabilities that lead to highly turbulent vortices that generate the pressure fluctuations responsible for sound. In order to reduce the noise associated with jet flow, the aerospace industry has focused on developing various technologies to disrupt shear layer turbulence and reduce the overall noise produced.
Turbofan engine noise propagates both upstream the inlet and downstream the primary nozzle and the by-pass duct. The main noise sources are the turbine and the compressor, the jet and the fan. The contribution of each noise source significantly evolved in the last decades: in typical 1960s design the jet was the main source whereas in modern turbofans the fan is the main noise source.
The fan noise is a tonal noise and its signature depends on the fan rotational speed:at low speed, the fan noise is due to the interaction of the blades with the distorted flow injected in the engine; this happens for example during the approach;
at high engine ratings, the fan tip is supersonic and this allows intense rotor-locked duct modes to propagate upstream; this noise is known as "buzz saw" and is typical at take-off.
All modern turbofan engines are equipped with acoustic liners to damp the noise generated. These are installed in the nacelle, and they extend as much as possible to cover the largest area. The acoustic performance of the engine can be experimentally evaluated by means of ground tests or in dedicated experimental test rigs.
In the aerospace industry, chevrons are the saw tooth patterns on the trailing edges of some jet engine nozzles that are used for noise reduction. Their principle of operation is that, as hot air from the engine core mixes with cooler air blowing through the engine fan, the shaped edges serve to smooth the mixing, which reduces noise-creating turbulence. Chevrons were developed with the help of NASA. Some notable examples of such engines include GEnx and Rolls-Royce Trent 1000.
The turbine blades in a turbofan engine are subject to high heat and stress, and require special fabrication. New material construction methods and material science have allowed blades, which were originally polycrystalline (regular metal), to be made from lined up metallic crystals and more recently mono-crystalline (i.e., single crystal) blades, which can operate at higher temperatures with less distortion.
Nickel-based superalloys are used for HP turbine blades in almost all modern jet engines. The temperature capabilities of turbine blades have increased mainly through four approaches: the manufacturing (casting) process, cooling path design, thermal barrier coating (TBC), and alloy development.
Although turbine blade (and vane) materials have improved over the years, much of the increase in (HP) turbine inlet temperatures is due to improvements in blade/vane cooling technology. Relatively cool air is bled from the compression system, bypassing the combustion process, and enters the hollow blade or vane. The gas temperature can therefore be even higher than the melting temperature of the blade. After picking up heat from the blade/vane, the cooling air is dumped into the main gas stream. If the local gas temperatures are low enough, downstream blades/vanes are uncooled and not adversely affected.
Strictly speaking, cycle-wise the HP turbine rotor inlet temperature (after the temperature drop across the HPT stator) is more important than the (HP) turbine inlet temperature. Although some modern military and civil engines have peak RITs of the order of 1,560 °C (2,840 °F), such temperatures are only experienced for a short time (during take-off) on civil engines.
Aerodynamics is a mix of subsonic, transonic and supersonic airflow on a single fan/gas compressor blade in a modern turbofan. The airflow past the blades has to be maintained within close angular limits to keep the air flowing against an increasing pressure. Otherwise the air will come back out of the intake.
A 100 g turbine blade is subjected to 1,700 °C/3100 °F, at 17 bars/250 Psi and a centrifugal force of 40 kN/ 9,000 lbf, well above the point of plastic deformation and even above the melting point. Exotic alloys, sophisticated air cooling schemes and special mechanical design are needed to keep the physical stresses within the strength of the material. Rotating seals must withstand harsh conditions for 10 years, 20,000 missions and rotating at 10–20,000 rpm.
The Full Authority Digital Engine Control needs accurate data for controlling the engine. The critical turbine inlet temperature (TIT) is too harsh an environment, at 1,700 °C and 17 bars, for reliable sensors. During development of a new engine type a relation is established between a more easily measured temperature like Exhaust gas temperature and the TIT. The EGT is then used to make sure the engine doesn't run too hot.
The turbofan engine market is dominated by General Electric, Rolls-Royce plc and Pratt & Whitney, in order of market share. GE and SNECMA of France have a joint venture, CFM International. Pratt & Whitney also have a joint venture, International Aero Engines with Japanese Aero Engine Corporation and MTU of Germany, specializing in engines for the Airbus A320 family. Pratt & Whitney and General Electric have a joint venture, Engine Alliance selling a range of engines for aircraft such as the Airbus A380.
For airliners and cargo aircraft, the in-service fleet in 2016 is 60,000 engines and should grow to 103,000 in 2035 with 86,500 deliveries according to Flight Global. A majority will be medium-thrust engines for narrow-body aircraft with 54,000 deliveries, for a fleet growing from 28,500 to 61,000. High-thrust engines for wide-body aircraft, worth 40–45% of the market by value, will grow from 12,700 engines to over 21,000 with 18,500 deliveries. The regional jet engines below 20,000lb (89kN) fleet will grow from 7,500 to 9,000 and the fleet of turboprops for airliners will increase from 9,400 to 10,200. The manufacturers market share should be led by CFM with 44% followed by Pratt & Whitney with 29% and then Rolls-Royce and General Electric with 10% each.
GE Aviation, part of the General Electric Conglomerate, currently has the largest share of the turbofan engine market. Some of their engine models include the CF6 (available on the Boeing 767, Boeing 747, Airbus A330 and more), GE90 (only the Boeing 777) and GEnx (developed for the Boeing 747-8 & Boeing 787 Dreamliner and proposed for the Airbus A350, currently in development) engines. On the military side, GE engines power many U.S. military aircraft, including the F110, powering 80% of the US Air Force's F-16 Fighting Falcons, and the F404 and F414 engines, which power the Navy's F/A-18 Hornet and Super Hornet. Rolls-Royce and General Electric were jointly developing the F136 engine to power the Joint Strike Fighter, however, due to government budget cuts, the program has been eliminated.
Rolls-Royce plc is the second largest manufacturer of turbofans and is most noted for their RB211 and Trent series, as well as their joint venture engines for the Airbus A320 and McDonnell Douglas MD-90 families (IAE V2500 with Pratt & Whitney and others), the Panavia Tornado (Turbo-Union RB199) and the Boeing 717 (BR700). The Rolls-Royce AE 3007, developed by Allison Engine Company before its acquisition by Rolls-Royce, powers several Embraer regional jets. Rolls-Royce Trent 970s were the first engines to power the new Airbus A380. The famous thrust vectoring Pegasus - actually a Bristol Siddeley design taken on by Rolls-Royce when they took over that company - is the primary powerplant of the Harrier "Jump Jet" and its derivatives.
Pratt & Whitney is third behind GE and Rolls-Royce in market share. The JT9D has the distinction of being chosen by Boeing to power the original Boeing 747 "Jumbo jet". The PW4000 series is the successor to the JT9D, and powers some Airbus A310, Airbus A300, Boeing 747, Boeing 767, Boeing 777, Airbus A330 and MD-11 aircraft. The PW4000 is certified for 180-minute ETOPS when used in twinjets. The first family has a 94-inch (2.4 m) fan diameter and is designed to power the Boeing 767, Boeing 747, MD-11, and the Airbus A300. The second family is the 100 inch (2.5 m) fan engine developed specifically for the Airbus A330 twinjet, and the third family has a diameter of 112-inch (2.8 m) designed to power Boeing 777. The Pratt & Whitney F119 and its derivative, the F135, power the United States Air Force's F-22 Raptor and the international F-35 Lightning II, respectively. Rolls-Royce are responsible for the lift fan which will provide the F-35B variants with a STOVL capability. The F100 engine was first used on the F-15 Eagle and F-16 Fighting Falcon. Newer Eagles and Falcons also come with GE F110 as an option, and the two are in competition.
CFM International is a joint venture between GE Aircraft Engines and SNECMA of France. They have created the very successful CFM56 series, used on Boeing 737, Airbus A340, and Airbus A320 family aircraft.
Engine Alliance is a 50/50 joint venture between General Electric and Pratt & Whitney formed in August 1996 to develop, manufacture, sell, and support a family of modern technology aircraft engines for new high-capacity, long-range aircraft. The main application for such an engine, the GP7200, was originally the Boeing 747-500/600X projects, before these were cancelled owing to lack of demand from airlines. Instead, the GP7000 has been re-optimised for use on the Airbus A380 superjumbo. In that market it is competing with the Rolls-Royce Trent 900, the launch engine for the aircraft. The two variants are the GP7270 and the GP7277.
International Aero Engines is a Zürich-registered joint venture between Pratt & Whitney, MTU Aero Engines and Japanese Aero Engine Corporation. The collaboration produced the V2500, the second most successful commercial jet engine program in production today in terms of volume, and the third most successful commercial jet engine program in aviation history.
Williams International is a manufacturer of small gas turbine engines based in Walled Lake, Michigan, United States. It produces jet engines for cruise missiles and small jet-powered aircraft. They have been producing engines since the 1970s and the range produces between 1000 and 3600 pounds of thrust. The engines are used as original equipment on the Cessna CitationJet CJ1 through CJ4 and Cessna Mustang, Beechcraft 400XPR and Premier 1a and there are several development programs with other manufacturers. The range is also very popular with the re-engine market being used by Sierra Jet and Nextant to breathe new life into aging platforms.
Honeywell Aerospace is one of the largest manufacturer of aircraft engines and avionics, as well as a producer of auxiliary power units (APUs) and other aviation products. Headquartered in Phoenix, Arizona, it is a division of the Honeywell International conglomerate. Honeywell/ITEC F124 series is used in military jets, such as the Aero L-159 Alca and the Alenia Aermacchi M-346. The Honeywell HTF700 series is used in the Bombardier Challenger 300 and the Gulfstream G280. The ALF502 and LF507 turbofans are produced by a partnership between Honeywell and China's state-owned Industrial Development Corporation. The partnership is called the International Turbine Engine Co.
Aviadvigatel is a Russian manufacturer of aircraft engines that succeeded the Soviet Soloviev Design Bureau. The company currently offers several versions of the Aviadvigatel PS-90 engine that powers Ilyushin Il-96-300/400/400T, Tupolev Tu-204, Tu-214 series and the Ilyushin Il-76-MD-90. The company is also developing the new Aviadvigatel PD-14 engine for the new Russian MS-21 airliner.
Ivchenko-Progress is the Ukrainian aircraft engine company that succeeded the Soviet Ivchenko Design Bureau. Some of their engine models include Progress D-436 available on the Antonov An-72/74, Yakovlev Yak-42, Beriev Be-200, Antonov An-148 and Tupolev Tu-334 and Progress D-18T that powers two of the world's largest airplanes, Antonov An-124 and Antonov An-225.
NPO Saturn is a Russian aircraft engine manufacturer, formed from the mergers of Rybinsk and Lyul'ka-Saturn. Saturn's engines include Lyulka AL-31, Lyulka AL-41, NPO Saturn AL-55 and power many former Eastern Bloc aircraft, such as the Tupolev Tu-154. Saturn holds a 50% stake in the PowerJet joint venture with Snecma.
PowerJet is a 50-50 joint venture between Snecma (Safran) and NPO Saturn, created in July 2004. The company manufactures SaM146, the sole powerplant for the Sukhoi Superjet 100.
Klimov was formed in the early 1930s to produce and improve upon the liquid-cooled Hispano-Suiza 12Y V-12 piston engine for which the USSR had acquired a license. Currently, Klimov is the manufacturer of the Klimov RD-33 turbofan engines.
EuroJet Turbo GmbH is a multi-national consortium, the partner companies of which are Rolls Royce of the United Kingdom, Avio of Italy, ITP of Spain and MTU Aero Engines of Germany. Eurojet GmbH was formed in 1986 to manage the development, production, support, maintenance, support and sales of the EJ200 turbofan engine for the Eurofighter Typhoon.
Three Chinese corporations build turbofan engines. Some of these are licensed or reverse engineered versions of European and Russian turbofans, and the other are indigenous models, but all are in development phase. Shenyang Aircraft Corporation (manufacturer of Shenyang WS-10), Xi'an Aero-Engine Corporation (manufacturer of Xian WS-15) and Guizhou Aircraft Industry Corporation (manufacturer of Guizhou WS-13) manufacture turbofans.
Ishikawajima-Harima Heavy Industries is the Japan aircraft engine company. The company manufactures F3 for Kawasaki T-4, XF5-1 for ATD-X, F7 for Kawasaki P-1.
Gas Turbine Research Establishment is owned by DRDO of Government of India. It produced the GTRE GTX-35VS Kaveri turbofan intended to power HAL Tejas and HAL Advanced Medium Combat Aircraft being built by the Aeronautical Development Agency.
In the 1970s, Rolls-Royce/SNECMA tested a M45SD-02 turbofan fitted with variable pitch fan blades to improve handling at ultra low fan pressure ratios and to provide thrust reverse down to zero aircraft speed. The engine was aimed at ultra quiet STOL aircraft operating from city centre airports.
In a bid for increased efficiency with speed, a development of the turbofan and turboprop known as a propfan engine was created that had an unducted fan. The fan blades are situated outside of the duct, so that it appears like a turboprop with wide scimitar-like blades. Both General Electric and Pratt & Whitney/Allison demonstrated propfan engines in the 1980s. Excessive cabin noise and relatively cheap jet fuel prevented the engines being put into service.Afterburner
extra combustor immediately upstream of final nozzle (also called reheat)
afterburner on low-bypass turbofan engines.
Average stage loading
constant × (delta temperature)/[(blade speed) × (blade speed) × (number of stages)]
airstream that completely bypasses the core compression system, combustor and turbine system
bypass airflow /core compression inlet airflow
turbomachinery handling the airstream that passes through the combustor.
residual shaft power from ideal turbine expansion to ambient pressure after deducting core compression power
Core thermal efficiency
core power/power equivalent of fuel flow
afterburner (if fitted) not lit
exhaust gas temperature
engine pressure ratio
turbofan LP compressor
Fan pressure ratio
fan outlet total pressure/intake delivery total pressure
use of artificially high apparent air temperature to reduce engine wear
high-pressure compressor (also HPC)
Intake ram drag
penalty associated with jet engines picking up air from the atmosphere (conventional rocket motors do not have this drag term, because the oxidiser travels with the vehicle)
integrated engine pressure ratio
intermediate pressure compressor (also IPC)
intermediate pressure turbine (also IPT)
low-pressure compressor (also LPC)
low-pressure turbine (also LPT)
nozzle total gross thrust - intake ram drag (excluding nacelle drag, etc., this is the basic thrust acting on the airframe)
Overall pressure ratio
combustor inlet total pressure/intake delivery total pressure
thermal efficiency * propulsive efficiency
propulsive power/rate of production of propulsive kinetic energy (maximum propulsive efficiency occurs when jet velocity equals flight velocity, which implies zero net thrust!)
Specific fuel consumption (SFC)
total fuel flow/net thrust (proportional to flight velocity/overall thermal efficiency)
accelerating, marked by a delay
pressure of the fluid which is associated not with its motion but with its state
net thrust/intake airflow
rate of production of propulsive kinetic energy/fuel power
Total fuel flow
combustor (plus any afterburner) fuel flow rate (e.g., lb/s or g/s)
static pressure plus
kinetic energy term
Turbine rotor inlet temperature
gas absolute mean temperature at principal (e.g., HP) turbine rotor entry